A conventional gas turbine engine includes a compressor for providing compressed air to a combustor wherein it is mixed with fuel and ignited for generating combustion gases. The combustion gases are channeled firstly through a high pressure turbine and typically through a low pressure turbine disposed downstream therefrom for extracting energy to drive the compressor and provide output power in the form of a combustion gas exhaust jet or shaft power for rotating a fan for generating thrust for powering an aircraft in flight. The efficiency of the engine is directly related to the high pressure turbine inlet temperature of the combustion gases channeled thereto. Since the combustion gases are considerably hot, the turbine rotor blades are typically hollow and provided with conventional film cooling holes for providing effective cooling thereof for ensuring a useful operating life. A portion of the compressed air from the compressor is suitably channeled to the turbine blades for cooling, and since such compressed air does not undergo combustion in the combustor, the net efficiency of the engine is decreased.
Accordingly, turbine blades are continually being improved for reducing the required amount of compressor cooling air channeled thereto for improving overall efficiency of the engine while still providing acceptable life of the blade. For example, in advanced gas turbine engines being presently considered, substantial reductions in cooling airflow through the turbine blades is being considered along with a substantial increase in the turbine rotor inlet temperature for substantially improving the operating efficiency of the engine. This will require a substantial increase in heat transfer effectiveness in cooling the turbine blades using the reduced amount of cooling airflow.
A conventional turbine blade includes a concave pressure side and a convex suction side over which the static pressure of the combustion gases channeled thereover varies significantly. Compressed air is conventionally channeled upwardly through the blade dovetail and into conventional serpentine passages through the blade airfoil for the convective coding thereof. The several passages within the blade airfoil may have various heat transfer enhancement turbulator ribs or pins for increasing the heat transfer coefficient over that for a smooth wall. Furthermore, conventional film cooling holes are selectively provided around the surface of the airfoil as required for forming suitable film cooling air layers to protect the airfoil against the hot combustion gases. Since the leading edge of the airfoil is typically subjected to the highest heat flux from the combustion gases, it typically requires the greatest protection from the heat and the highest heat transfer enhancement from the compressed air being channeled through the airfoil.
In one conventional airfoil having a mid-chord radial flow channel, the compressed air is normally channeled therethrough with a fairly uniform parabolic radial velocity distribution with a maximum velocity near the center of the channel and lower velocities adjacent the pressure and suction side interior walls of the airfoil. However, it has been observed that by providing a plurality of film cooling holes along the radial channel to eject uniformly part of the inlet channel air flow out the airfoil external surface on only one side of the channel, such as the pressure side, with the other, suction side of the channel being imperforate, the radial velocity distribution of the compressed air through the channel is distorted with the peak velocity being moved transversely toward the film cooling holes. Correspondingly, the velocity of the compressed air adjacent the opposite, suction side decreases for a given average flow rate through the channel. Accordingly, the decreased compressed air velocity adjacent the imperforate wall results in a decrease of the convective heat transfer enhancement of that wall which typically requires an increase in the average velocity of the compressed air to offset the decrease at the wall to ensure effective cooling thereof.
Furthermore, the turbine blade must also be designed to provide an adequate backflow margin to ensure that the combustion gases are not allowed to backflow through the film cooling holes into the blade airfoil during operation. The pressure of the compressed air inside the airfoil is, therefore, predeterminedly selected to be suitably larger than the pressure of the combustion gases flowing over the airfoil to ensure the forward flow of the compressed air from the interior of the airfoil through the film cooling holes to the exterior of the airfoil. Since the leading edge region of a typical high pressure turbine blade includes film cooling holes along both the pressure and suction sides of the airfoil, the pressure of the compressed air inside the airfoil must be suitably large to provide an effective backflow margin through the pressure side film cooling holes adjacent the leading edge which are subject to the highest pressure from the combustion gases flowable over the airfoil. However, since the pressure of the combustion gases channeled over the suction side of the airfoil is necessarily lower than that over the pressure side, the pressure ratio of the compressed air inside the airfoil relative to the suction side adjacent the leading edge is relatively high, and higher than the pressure ratio across the pressure side at the leading edge, which increases the ejection velocity of the film cooling air through the suction side leading edge film cooling holes. This may lead to a condition known as blowoff wherein the film cooling air initially breaks free from the airfoil suction side as it is ejected from the film cooling holes before reattaching to the suction side downstream therefrom. This leads to a decrease in the air film effectiveness and cooling capability of the film cooling air in this region.
In both of the above situations, the internal heat transfer coefficient and the film cooling effectiveness of the compressed air channeled through the airfoil are decreased, which requires even more air, for example, to ensure acceptable cooling of the turbine blade, which decreases overall efficiency. Furthermore, it is desirable to channel the compressed air through the airfoil with as little pressure losses therein as possible to further increase the overall efficiency of the engine.
Since a high pressure turbine blade typically includes several serpentine flow passages therein and film cooling holes along the pressure and suction sides thereof, as well as discharge holes through the trailing edge thereof, the blade must be suitably tested to ensure that suitable flow is obtained through all of the holes. In a turbine blade having two or more independent cooling air channels or circuits therethrough, for example one circuit channeling cooling air from the mid-chord and out through the leading edge holes, and a second circuit channeling air through the mid-chord and out the trailing edge holes, the several holes of each circuit must be independently plugged, for example, to test the operation of the other circuit. This is a relatively complex procedure required for ensuring effective operation of all the apertures through the blade airfoil.